GAS TURBINES P1

The turbine rotation is caused by the work produced by the fluid while expanding from the high pressure at the combustor discharge to ambient air pressure at the turbine exhaust.. Gas tu

57.1 INTRODUCTION

57.1.1 Basic Operating Principles

Gas turbines are heat engines based on the Brayton thermodynamic cycle This cycle is one of the four that account for most of the heat engines in use Other cycles are the Otto, Diesel and Rankine The Otto and Diesel cycles are cyclic in regard to energy content Steady-flow, continuous energy transfer cycles are the Brayton (gas turbine) and Rankine (steam turbine) cycles The Rankine cycle involves condensing and boiling of the working fluid, steam, and utilizes a boiler to transfer heat to the working fluid The working fluid in the other cycles is generally air, or air plus combustion products The Otto, Diesel and Brayton cycles are usually internal combustion cycles wherein the fuel is burned in the working fluid In summary, the Brayton cycle is differentiated from the Otto and Diesel cycle in that it is continuous, and from the Rankine in that it relies on internal combustion, and does not involve a phase change in the working fluid

In all cycles, the working fluid experiences induction, compression, heating, expansion, and ex-haust In a non-steady cycle, these processes are performed in sequence in the same closed space,

This chapter was written as an update to chapter 72 of the Handbook's previous edition Much of the structure and significant portions of the text of the previous edition's chapter is retained The new edition has increased emphasis on the most significant current and future projected gas turbine con-figurations and applications Thermodynamic cycle variations are presented here in a consistent for-mat, and the description of current cycles replaces the discussions of some interesting and historical, but less significant, cycles described in the earlier edition In addition, there is a new discussion of economic and regulatory trends, of supporting technologies, and their interconnection with gas turbine development The author of the previous version had captured the history of the gas turbine's de-velopment, and this history is repeated and supplemented here

Mechanical Engineers' Handbook, 2nd ed., Edited by Myer Kutz.

ISBN 0-471-13007-9 © 1998 John Wiley & Sons, Inc

CHAPTER 57

GAS TURBINES

Harold Miller

GE Power Systems

Schenectady, New York

57.1 INTRODUCTION 1723

57.1.1 Basic Operating

Principles 1723

57 1.2 A Brief History of Gas

Turbine Development

and Use 1727

57.1.3 Components,

Characteristics and

Capabilities 1728

57.1.4 Controls and Accessories 1737

57.1.5 Gas Turbine Operation 1740

57.2 GAS TURBINE

PERFORMANCE 1740

57.2.1 Gas Turbine

Configurations and

Cycle Characteristics 1740

57.2.2 Trends in Gas Turbine

Design and Performance 1747

57.3 APPLICATIONS 1749 57.3.1 Use of Exhaust Heat in Industrial Gas Turbines 1749 57.3.2 Integrated Gasification

Combined Cycle 1751 57.3.3 Applications in Electricity Generation 1753 57.3.4 Engines for Aircraft 1755 57.3.5 Engines for Surface

Transportation 1757 57.4 EVALUATIONAND

SELECTION 1759 57.4.1 Maintenance Intervals,

Availability, and Reliability 1759 57.4.2 Selection of Engine and System 1761

formed by a piston and cylinder that operate on the working fluid one mass at a time In contrast, the working fluid flows through a steam turbine power plant or gas turbine engine, without interrup-tion, passing continuously from one single-purpose device to the next

Gas turbines are used to power aircraft and land vehicles, to drive generators (alternators) to produce electric power, and to drive other devices, such as pumps and compressors Gas turbines in production range in output from below 50 kW to over 200 MW Design philosophies and engine configurations vary significantly across the industry Aircraft engines are optimized for high power-to-weight ratios, while heavy-duty, industrial, and utility gas turbines are heavier, being designed for low cost and long life in severe environments

The arrangement of a simple gas turbine engine is shown in Fig 57.1a The rotating compressor acts to raise the pressure of the working fluid and force it into the combustor The turbine rotation

is caused by the work produced by the fluid while expanding from the high pressure at the combustor discharge to ambient air pressure at the turbine exhaust The resulting mechanical work drives the mechanically connected compressor and output load device

The nomenclature of the gas turbine is not standardized In this chapter, the term blading refers

to all rotating and stationary airfoils in the gas path Turbine (expander) section rotating blades are

buckets, a term derived from steam turbine practice Turbine section stationary blades are nozzles The combustion components in contact with the working fluid are called combustors; major

com-bustor components are fuel nozzles and combustion liners Some comcom-bustors (Can-annular and silo-types) have transition pieces that conduct hot gas from the combustion liners to the first-stage nozzles

A stage of the compressor consists of a row of rotor blades, all at one axial position in the gas turbine, and the stationary blade row downstream of it A turbine stage consists of a set of nozzles occupying one axial location and the set of buckets immediately downstream Rotating blading is attached either to a monolithic rotor structure or to individual discs or wheels designed to support

the blading against centrifugal force and the aerodynamic loads of the working fluid The terms discs and wheels are used interchangeably.

Gas turbine performance is established by three basic parameters: mass flow, pressure ratio, and firing temperature Compressor, combustor, and turbine efficiency have significant, but secondary, effects on performance, as do inlet and exhaust systems, turbine gas path and rotor cooling, and heat loss through turbine and combustor casings

In gas turbine catalogues and other descriptive literature, mass flow is usually quoted as com-pressor inlet flow, although turbine exit flow is sometimes quoted Output is proportional to mass flow

Pressure ratio is quoted as the compressor pressure ratio Aircraft engine practice is to define the ratio as the total pressure at the exit of the compressor blading divided by the total pressure at the inlet of the compressor blading Industrial/utility turbine manufacturers generally refer to the static pressure in the plenum downstream of the compressor discharge diffuser (upstream of the combustor) divided by the total pressure downstream of the inlet filter and upstream of the inlet of the gas turbine Similarly, there are various possibilities for defining turbine pressure ratio All definitions yield values within 1 or 2% of one another Pressure ratio is the primary determinant of simple cycle gas turbine efficiency High pressure results in high simple cycle efficiency

Firing temperature is defined differently by each manufacturer, and the differences are significant Heavy-duty gas turbine manufacturers use three definitions There is an ISO definition of firing temperature, which is a calculated temperature The compressor discharge temperature is increased

by a calculated enthalpy rise based on the compressor inlet air flow and the fuel flow This definition

is valuable in that it can be used to compare gas turbines or monitor changes in performance through calculations made on the basis of field measurements To determine ISO firing temperature, one does not require knowledge of the secondary flows within the gas turbine A widely used definition of

Fig 57.1 Simple engine type: (a) open cycle; (b) closed cycle (diagrammatic).1

firing temperature is the average total temperature in the exit plane of the first stage nozzle This definition is used by General Electric for its industrial engines Westinghouse refers to "turbine inlet temperature," the temperature of the gas entering the first stage nozzle Turbine inlet temperature is approximately 10O0C above nozzle exit firing temperature, which is in turn approximately 10O0C above ISO firing temperature Since firing temperature is commonly used to compare the technology level of competing gas turbines, it is important to compare on one definition of this parameter Aircraft engines and aircraft-derivative industrial gas turbines have other definitions One nomen-clature establishes numerical stations—here, station 3.9 is combustor exit and station 4.0 is first-stage nozzle exit Thus, T39 is very close to "turbine inlet temperature" and T40 is approximately equal

to GE's "firing temperature." There are some subtle differences relating to the treatment of the leakage flows near the first-stage nozzle This nomenclature is based on SAE ARP 755A, a recom-mended practice for turbine engine notation

Firing temperature is a primary determiner of power density (specific work) and combined cycle (Brayton-Rankine) efficiency High firing temperature increases the power produced by a gas turbine

of a given physical size and mass flow The pursuit of higher firing temperatures by all manufacturers

of large, heavy-duty gas turbines used for electrical power generation is driven by the economics of high combined cycle efficiency

Pressures and temperatures used in the following descriptions of gas turbine performance will be total pressures and temperatures Absolute, stagnation, or total values are those measured by instru-ments that face into the approaching flow to give an indication of the energy in the fluid at any point The work done in compression or expansion is proportional to the change of stagnation temperature

in the working fluid, in the form of heating during a compression process or cooling during an expansion process The temperature ratio, between the temperatures before and after the process, is

related to the pressure ratio across the process by the expression T b IT a = (P b /P a } (y ~ l)/y , where y is

the ratio of working fluid specific heats at constant pressure and volume The temperature and pressure are stagnation values It is the interaction between the temperature change and ratio, at different starting temperature levels, that permits the engine to generate a useful work output

This relationship between temperature and pressure can be demonstrated by a simple numerical example using the Kelvin scale for temperature For a starting temperature of 30O0K (270C), a tem-perature ratio of 1.5 yields a final temtem-perature of 45O0K and a change of 15O0C Starting instead at 40O0K, the same ratio would yield a change of 20O0C and a final temperature of 60O0K The equivalent

pressure ratio would ideally be 4.13, as calculated from solving the preceding equation for P b IP a \

P b /P a = T b /Tl /y ~ l = 1.51-4'0-4 = 4.13 These numbers show that, working over the same temperature ratio, the temperature change and, therefore, the work involved in the process vary in proportion to the starting temperature level.2

This conclusion can be depicted graphically If the temperature changes are drawn as vertical

lines a-b and c-d, and are separated horizontally to avoid overlap, the resultant is Fig 57.2a

As-suming the starting and finishing pressures to be the same for the two processes, the thin lines through

a-d and b-c depict two of a family of lines of constant pressure, which diverge as shown In this

ideal case, expansion processes could be represented by the same diagram, simply by proceeding

down the lines b-a and c-d Alternatively, if a-b is taken as a compression process, b-c as heat addition, c-d as an expansion process, and d-a as a heat rejection process, then the figure a-b-c-d-a

represents the ideal cycle to which the working fluid of the engine is subjected

Over the small temperature range of this example, the assumption of constant gas properties is justified In practice, the 3270C (60O0K) level at point d is too low a temperature from which to start

Fig 57.2 Temperature changes and temperature-entropy diagram for ideal

simple gas turbine cycles

the expansion Figure 57.2b is more realistic Here, the lines of constant pressure have been

con-structed for ideal gas-air properties that are dependent upon temperature Expansion begins from a temperature of 125O0C With a pressure ratio of 16:1, the end point of the expansion is approximately 48O0C Now a-b represents the work input required by the compressor Of the expansion work capacity c-d, only the fraction c-d' is required to drive the compressor An optical illusion makes

it appear otherwise, but line a-d' is displaced vertically from line b-c by the same distance

every-where The remaining 4350C, line d'-d, is energy that can be used to perform useful external work,

by further expansion through the turbine or by blowing through a nozzle to provide jet thrust

Now consider line b-c The length of its vertical projection is proportional to the heat added The

ability of the engine to generate a useful output arises from its use of the energy in the input fuel flow, but not all of the fuel energy can be recovered usefully In this example, the heat input pro-portional to 1250-350 = 90O0C compares with the excess output proportional to 4350C (line d'-d}

to represent an efficiency of (435/900), or 48% If more fuel could be used, raising the maximum temperature level at the same pressure, then more useful work could be obtained at nearly the same efficiency

The line d-a represents heat rejection This could involve passing the exhaust gas through a

cooler before returning it to the compressor, and this would be a closed cycle But, almost universally,

d-a reflects discharge to the ambient conditions and intake of ambient air (Fig 57.1Z?) Figure 57.Ia

shows an open-cycle engine, which takes air from the atmosphere and exhausts back to the

atmos-phere In this case, line d-a still represents heat rejection, but the path from d to a involves the whole atmosphere and very little of the gas finds its way immediately from e to a It is fundamental to this

cycle that the remaining 4650C, the vertical projection of line d-a, is wasted heat because point d is

at atmospheric pressure The gas is therefore unable to expand further and so can do no more work Designers of simple cycle gas turbines—including aircraft engines—have pursued a course of reducing exhaust temperature through increasing cycle pressure ratio, which improves the overall

efficiency Figure 57.3 is identical to Fig 51.2b except for the pressure ratio, which has been increased

from 16:1 to 24:1 The efficiency is calculated in the same manner The total turbine work is proportional to the temperature difference across the turbine, 1250-410 = 84O0C The compressor work, proportional to 430-15 = 4150C, is subtracted from the turbine temperature drop 840-415 =

4250C The heat added to the cycle is proportional to 1250-430 = 82O0C The ratio of the net work

to the heat added is 425/820 = 52% The approximately 8% improvement in efficiency is accom-panied by a 7O0C drop in exhaust temperature When no use is made of the exhaust heat, the 8% efficiency may justify the mechanical complexity associated with higher pressure ratios Where there

is value to the exhaust heat, as there is in combined Brayton-Rankine cycle power plants, the lower pressure ratio may be superior Manufacturers forecast their customer requirements and understand

Fig 57.3 Simple cycle gas turbine temperature-entropy diagram for high (24:1) pressure ratio

and 125OC firing temperature

the costs associated with cycle changes and endeavor to produce gas turbines featuring the most economical thermodynamic designs

The efficiency levels calculated in the preceding example are very high because many factors have been ignored for the sake of simplicity Inefficiency of the compressor increases the compressor work demand, while turbine inefficiency reduces turbine work output, thereby reducing the useful work output and efficiency The effect of inefficiency is that, for a given temperature change, the compressor generates less than the ideal pressure level while the turbine expands to a higher tem-perature for the same pressure ratio There are also pressure losses in the heat addition and heat rejection processes There may be variations in the fluid mass flow rate and its specific heat (energy input divided by consequent temperature rise) around the cycle These factors can easily combine to reduce the overall efficiency

57.1.2 A Brief History of Gas Turbine Development and Use

The use of a turbine driven by the rising flue gases above a fire dates back to Hero of Alexandria in

150 BC It was not until AD 1791 that John Barber patented the forerunner of the gas turbine, proposing the use of a reciprocating compressor, a combustion system, and an impulse turbine Even then, he foresaw the need to cool the turbine blades, for which he proposed water injection The year 1808 saw the introduction of the first explosion type of gas turbine, which in later forms used valves at entry and exit from the combustion chamber to provide intermittent combustion in a closed space The pressure thus generated blew the gas through a nozzle to drive an impulse turbine These operated successfully but inefficiently for Karavodine and Holzwarth from 1906 onward, and the type died out after a Brown, Boveri model was designed in 1939.3

Developments of the continuous flow machine suffered from lack of knowledge, as different configurations were tried Stolze in 1872 designed an engine with a seven-stage axial flow compressor, heat addition through a heat exchanger by external combustion, and a 10-stage reaction turbine It was tested from 1900 to 1904 but did not work because of its very inefficient compressor Parsons was equally unsuccessful in 1884, when he tried to run a reaction turbine in reverse as a compressor These failures resulted from the lack of understanding of aerodynamics prior to the advent of aircraft

As a comparison, in typical modern practice, a single-stage turbine drives about six or seven stages

of axial compressor with the same mass flow

The first successful dynamic compressor was Rateau's centrifugal type in 1905 Three assemblies

of these, with a total of 25 impellers in series giving an overall pressure ratio of 4, were made by Brown, Boveri and used in the first working gas turbine engine, built by Armengaud and Lemale in the same year The exhaust gas heated a boiler behind the turbine to generate low-pressure steam, which was directed through turbines to cool the blades and augment the power Low component efficiencies and flame temperature (8280K) resulted in low work output and an overall efficiency of 3% By 1939, the use of industrial gas turbines had become well established: experience with the Velox boiler led Brown, Boveri into diverging applications; a Hungarian engine (Jendrassik) with axial flow compressor and turbine used regeneration to achieve an efficiency of 0.21; and the Sun Oil Co in the United States was using a gas turbine engine to improve a chemical process.2

The history of gas turbine engines for aircraft propulsion dates from 1930, when Frank Whittle saw that its exhaust gas conditions ideally matched the requirements for jet propulsion and took out

a patent.4 His first model was built by British Thomson-Houston and ran as the Power Jets Type U

in 1937, with a double-sided centrifugal compressor, a long combustion chamber that was curled round the outside of the turbine and an exhaust nozzle just behind the turbine Problems of low compressor and turbine efficiency were matched by hardware problems and the struggle to control the combustion in a very small space Reverse-flow, can-annular combustors were introduced in 1938, the aim still being to keep the compressor and turbine as close together as possible to avoid shaft whirl problems (Fig 57.4) Whittle's first flying engine was the Wl, with 850 Ib thrust, in 1941 It was made by Rover, whose gas turbine establishment was taken over by Rolls-Royce in 1943 A General Electric version of the Wl flew in 1941 A parallel effort at General Electric led to the development of a successful axial-flow compressor This was incorporated in the first turboprop engine, the TGlOO, later designated the T31 This engine, first tested in May of 1943, produced 1200 horsepower from an engine weighing under 400 kg Flight testing followed in 1949 An axial-compressor turbojet version was also constructed, designated the J35 It flew in 1946 The axial-compressor

of this engine evolved to the compressor of the GE MS3002 industrial engine, which was introduced

in 1950 and is still in production.5

A Heinkel experimental engine flew in Germany in 1939 Several jet engines were operational

by the end of the Second World War, but the first commercial engine did not enter service until 1953, the Rolls-Royce Dart turboprop in the Viscount, followed by the turbojet de Havilland Ghost in the Comet of 1954 The subsequent growth of the use of jet engines has been visible to most of the world, and has forced the growth of design and manufacturing technology.6 By 1970, a range of standard configurations for different tasks had become established, and some aircraft engines were established in industrial applications and in ships

Gas turbines entered the surface transportation fields also during their early stages of development The first railway locomotive application was in Switzerland in 1941, with a 2200-hp Brown, Boveri

Fig 57.4 Simplified arrangement of an early Whittle jet engine, with double-sided compressor

and reverse-flow combustion chambers (Redrawn from Ref 4 by permission of the Council of

the Institution of Mechanical Engineers.)

engine driving an electric generator and electric motors driving the wheels The engine efficiency approached 19%, using regeneration The next decade saw several similar applications of gas turbines

by some 43 different manufacturers A successful application of gas turbines to transportation was the 4500 draw-bar horsepower engine, based on the J35 compressor Twenty-five locomotives so equipped were delivered to the Union Pacific railroad between 1952 and 1954 The most powerful locomotive gas turbine was the 8500-hp unit offered by General Electric to the Union Pacific railroad for long-distance freight service.7 This became the basis of the MS5001 gas turbine, which is the most common heavy-duty gas turbine in use today Railroad applications continue today, but relying

on a significantly different system Japan Railway uses large stationary gas turbines to generate power transmitted by overhead lines to their locomotives

Automobile and road vehicle use started with a Rover car of 1950, followed by Chrysler and other companies, but commercial use has been limited to trucks, particularly by Ford Automotive gas turbine development has been largely independent of other types, and has forced the pace of development of regenerators

57.1.3 Component Characteristics and Capabilities

Compressors

Compressors used in gas turbines are of the dynamic type, wherein air is continuously ingested and raised to the required pressure level—usually, but not necessarily, between 8 and 40 atmospheres Larger gas turbines use axial types; smaller ones use radial outflow centrifugal compressors Some smaller gas turbines use both—an axial flow compressor upstream of a centrifugal stage

Axial compressors feature an annular flowpath, larger in cross-section area at the inlet than at the discharge Multiple stages of blades alternately accelerate the flow of air and allow it to expand, recovering the dynamic component and increasing pressure Both rotating and stationary stages con-sist of cascades of airfoils, as can be seen in Fig 57.5 Physical characteristics of the compressor determine many aspects of the gas turbine's performance Inlet annulus area establishes the mass flow of the gas turbine Rotor speed and mean blade diameter are interrelated, since optimum blade velocities exist A wide range of pressure ratios can be provided, but today's machines feature com-pressions from 8:1 to as high as 40:1 The higher pressure ratios are achieved using two compressors operating in series at different rotational speeds The number of stages required is partially dependent

on the pressure ratio required, but also on the sophistication of the blade aerodynamic design that is applied Generally, the length of the compressor is a function of pressure ratio, regardless of the number of stages Older designs have stage pressure ratios of 1.15:1 Low-aspect ratio blading designed with three-dimensional analytical techniques have stage pressure ratios of 1.3:1 There is

a trend toward fewer stages of blades of more complicated configuration Modern manufacturing techniques make more complicated forms more practical to produce, and minimizing parts count usually reduces cost

Centrifugal compressors are usually chosen for machines of below 2 or 3 MW in output, where their inherent simplicity and ruggedness can largely offset their lower compression efficiency Such compressors feature a monolithic rotor with a shaped passage leading from the inlet circle or annulus

to a volute at the outer radius, where the compressed air is collected and directed to the combustor The stator contains no blades or passages and simply provides a boundary to the flow path, three sides of which are machined or cast into the rotor Two or more rotors can be used in series to achieve the desired pressure ratio within the mechanical factors that limit rotor diameter at a given rotational speed

Fig 57.5 Diagram, and photos of centrifugal compressor rotor (courtesy of Nuovo Pignone

Company) and axial compressor during assembly (courtesy of General Electric Company).

Two efficiency definitions are used to describe compressor performance Polytropic efficiency characterizes the aerodynamic efficiency of low-pressure-ratio individual stages of the compressor Isentropic, or adiabatic, efficiency describes the efficiency of the first step of the thermodynamic

process shown in Fig 57.6 (the path from a to b) The isentropic efficiency can be calculated from

the temperatures shown for the compression process on this figure The isentropic temperature rise

is for the line a-b: 3350C The actual rise is shown by line a-b', and this rise is 3720C The compressor efficiency T\ C is the ratio 335/372 = 90%

Successful compressor designs achieve high component efficiency while avoiding compressor surge or stall—the same phenomenon experienced when airplane wings are forced to operate at too high an angle of attack at too low a velocity Furthermore, blade and rotor structures must be designed

to avoid vibration problems These problems occur when natural frequencies of components and assemblies are coincident with mechanical and aerodynamic stimuli, such as those encountered as blades pass through wakes of upstream blades The stall phenomenon may occur locally in the compressor or even generally, whereupon normal flow through the machine is disrupted A com-pressor must have good stall characteristics in order to operate at all ambient pressures and temper-atures and to operate through the start, acceleration, load, load-change, unload, and shutdown phases

of turbine operation Compressors are designed with features and mechanisms for avoiding stall These include air bleed at various points, variable-angle stator (as opposed to rotor) blades, and multiple spools

Recent developments in the field of computational fluid dynamics (CFD) provide analytical tools that allow designers to substantially reduce aerodynamic losses due to shock waves in the supersonic flow regions Using this technique, stages that have high tip Mach numbers can attain efficiencies comparable to those of completely subsonic designs With these tools, compressors can

be designed with higher tip diameters, hence higher flows The same tools permit the design of low

Fig 57.6 Temperature-entropy diagram showing the effect of compressor

and turbine efficiency

aspect ratio, high stage pressure ratio blades for reducing the number of blade rows Both capabilities contribute to lower cost gas turbine designs with no sacrifice in performance

Gas Turbine Combustion System

The gas turbine combustor is a device for mixing large quantities of fuel and air and burning the resulting mixture A flame burns best when there is just enough fuel to react with the available

oxygen This is called a stoichiometric condition, and combustion here produces the fastest chemical

reaction and the highest flame temperatures, compared with excess air (fuel-lean) and excess fuel

(fuel-rich) conditions, where reaction rates and temperatures are lower The term equivalence ratio

is used to describe the ratio of fuel to air relative to the stoichiometric condition An equivalence ratio of 1.0 corresponds to the stoichiometric condition Under fuel-lean conditions, the ratio is less than 1, and under fuel-rich conditions it is greater than 1 The European practice is to use the reciprocal, which is the Lambda value (X)

In a gas turbine, since air is extracted from the compressor for cooling the combustor, buckets, nozzles, and other components and to dilute the flame—as well as support combustion—the overall equivalence ratio is far less than the value in the flame zone, ranging from 0.4-0.5 (X = 2.5 to 2).9

Historically, the design of combustors required providing for the near-stoichiometric mixture of fuel and air locally The combustion in this near-stoichiometric situation results in a diffusion flame

of high temperature Near-stoichiometric conditions produce a stable combustion front without re-quiring designers to provide significant flame-stabilizing features Since the temperatures generated

by the burning of a stoichiometric mixture greatly exceed those at which materials are structurally sound, combustors have to be cooled, and also the gas heated by the diffusion flame must be cooled

by dilution before it becomes the working fluid of the turbine

Gas turbine operation involves a startup cycle that features ignition of fuel at 20% of rated operating speed where air flow is proportionally lower Loading, unloading, and part-load operation, however, require low fuel flow at full compressor speed, which means full air flow Thermodynamic cycles are such that the lowest fuel flow per unit mass flow of air through the turbine exists at full speed and no-load The fuel flow here is about 1/6 of the full-load fuel flow Hence, the combustion system must be designed to operate over a 6:1 range of fuel flows with full rated air flow Manufacturers have differed on gas turbine combustor construction in significant ways Three basic configurations have been used: annular, can-annular, and "silo" combustors All have been used successfully in machines with firing temperatures up to UOO0C Annular and can-annular combustors feature a combustion zone uniformly arranged about the centerline of the engine All aircraft engines and most industrial gas turbine feature this type of design A significant number of units equipped with silo combustors have been built as well Here, one or two large combustion vessels are

con-structed on top of or beside the gas turbine All manufacturers of large machines have now abandoned silo combustors in their state-of-the-art products The can-annular, multiple combustion chamber assembly consists of an arrangement of cylindrical combustors, each with a fuel injection system, and a transition piece that provides a flow path for the hot gas from the combustor to the inlet of the turbine Annular combustors have fuel nozzles at their upstream end and an inner and outer liner surface extending from the fuel nozzles to the entrance of the first-stage stationary blading No transition piece is needed

The current challenge to combustion designers is providing the cycle with a sufficiently high firing temperature while simultaneously limiting the production of oxides of nitrogen, NOx, which refers to NO and NO2 Very low levels of NOx have been achieved in special low-emission combus-tors NOx is formed from the nitrogen and oxygen in the air when it is heated The nitrogen and oxygen combine at a significant rate at temperatures above 150O0C, and the formation rate increases

exponentially as temperature increases Even with the high gas velocities in gas turbines, NO x emis-sions will reach 200 parts per million by volume, dry (ppmvd), in gas turbines with conventional

combustors and no NO x abatement features Emissions standards throughout the world vary, but many

parts of the world require gas turbines to be equipped to control NO x to below 25 parts per million

by volume, dry (ppmvd) at base load

Emissions

Combustion of common fuels necessarily results in the emission of water vapor and carbon dioxide Combustion of near-stoichiometric mixtures results in very high temperatures Oxides of nitrogen are formed as the oxygen and nitrogen in the air combine, and this happens at gas turbine combustion temperatures Carbon monoxide forms when the combustion process is incomplete Unburned hydro-carbons (UHC) are discharged as well when combustion is incomplete Other pollutants are attributed

to fuel; principal among these is sulfur Gas turbines neither add nor remove sulfur; hence, what sulfur enters the gas turbine in the fuel exits as SO2 in the exhaust

Much of the gas turbine combustion research and development of the 1980s and 1990s focused

on lowering NO x production in mechanically reliable combustors while maintaining low CO and

UHC emissions Early methods of reducing NO x emissions included removing it from the exhaust

by selective catalytic reduction (SCR) and by diluent injection, that is, the injection of water or steam into the combustor These methods continue to be employed The lean-premix combustors now in general use are products of ongoing research

Thermal NO x is generally regarded as being generated by a chemical reaction sequence called the Zeldovich mechanism,10 and the rate of NC x formation is proportional to temperature, as shown in Fig 57.7 In practical terms, a conventional gas turbine emits approximately 200 ppmvd when its

combustors are not designed to control NO x This is because a significant portion of the combustion

zone has stoichiometric or near-stoichiometric conditions, and temperatures are high Additional oxygen, and of course nitrogen on the boundary of the flame, is heated to sufficiently high

temper-atures, and held at these temperatures for sufficient time, to produce NO x

Water- and steam-injected combustors achieve low flame temperatures by placing diluent in the

neighborhood of the reacting fuel and air Among low NO x combustion systems operating today, water and steam injection is the most common means of flame temperature reduction Several hundred large industrial turbines operating with steam or water injection have accumulated over 2-1/2 million hours of service Water is not the only diluent used for NOx control In the case of integrated gasification combined cycle plants, nitrogen and CO2 are available and can be introduced into the combustion region The NOx emissions measured at the Cool Water IGCC plant in the United States rival those of the cleanest natural gas plants in the world.11

Water or steam injection can achieve levels that satisfy all current standards, but water consump-tion is sometimes not acceptable to the operator because of cost, availability, or the impact on efficiency Steam injection sufficient to reduce NOx emissions to 25 ppmvd can increase fuel con-sumption in combined cycle power plants by over 3% Water injection increases fuel use by over 4% for the same emissions level In base-load power plants, fuel cost is so significant that it has caused the development of systems that do not require water.12

In all combustion processes, when a molecule of methane combines with two molecules of ox-ygen, a known and fixed amount of heat is released When only these three molecules are present,

a minimum amount of mass is present to absorb the energy not radiated and the maximum temperature

is realized Add to the neighborhood of the reaction the nitrogen as found in air (four times the volume of oxygen involved in the reaction) and the equilibrium temperature is lower When even more air is added to the combustion region, more mass is available to absorb the energy and the resulting observable temperature is lower still The same can be achieved through the use of excess fuel Thus, moving away from the stoichiometric mixture means that observable flame temperature

is lowered and the production of NOx is also reduced On a microscopic level, lean-burning low-NOx

combustors are designed to force the chemical reaction to take place in such a way that the energy released is in the neighborhood of as much mass not taking part in the reaction as possible By transferring heat to neighboring material immediately, the time-at-temperature is reduced On a larger

Equivalence Ratio Fig 57.7 NOx formation rate driven by temperature (drawn from figure in Ref 9; courtesy of

General Electric Company)

scale, a high measurable temperature will never be reached in a well-mixed lean system and thus

NOx generation is minimized Both rich-mixture and lean-mixture systems have led to low NOx

schemes Although those featuring rich flames followed by lean burning zones are sometimes sug-gested for situations where there is nitrogen in the fuel, most of today's systems are based on lean burning

Early lean premix dry low-NOx combustors were operated in GE gas turbines at the Houston Light and Power Wharton Station in 1980 in the United States, in Mitsubishi units in Japan in 1983, and were introduced in Europe in 1986 by Siemens KWU These combustors control the formation

of NOx by premixing fuel with air prior to its ignition while conventional combustors mix essentially

at the instant of ignition Dry low-NOx combustors, as the name implies, achieve NOx control without consuming water and without imposing efficiency penalties on combined-cycle plants

Figures 57.8 and 57.9 show dry low-NOx combustors developed for large gas turbines In the GE system, several premixing chambers are located at the head end of the combustor A fuel nozzle assembly is located in the center of each chamber By the manipulation of valves external to the gas turbine, fuel can be directed to several combinations of chambers and to various parts of the fuel nozzles This is to permit the initial ignition of the fuel and to maintain a relatively constant local fuel-air ratio at all load levels There is one flame zone, immediately downstream of the premixing chambers The Westinghouse combustor illustrated in Fig 57.9 has three concentric premixing cham-bers The two nearest the centerline of the combustor are designed to swirl the air passing through them in opposite directions and discharge into the primary combustion zone The third, which has a longer passage, is directed to the secondary zone Modulating fuel flow to the various mixing passages and combustion zones ensures low NOx production over a wide range of operating temperatures Both the combustors shown are designed for state-of-the-art, high-firing-temperature gas turbines LoW-NOx combustors feature multiple premixing features and a more complex control system than more conventional combustors, to achieve stable operation over the required range of operating conditions The reason for this complexity is explained with the aid of Fig 57.10 Conventional combustors operate with stability over a wide range of fuel-air mixtures—between the rich and lean flammability limits A sufficiently wide range of fuel flows could be burned in a combustor with a fixed air flow, to match the range of load requirements from no-load to full-load In a low-NOx

combustor, the fuel-air mixture feeding the flame must be regulated between the point of flame loss and the point where the NOx limit is exceeded When low gas turbine output is required, the air premixed with the fuel must be reduced to match the fuel flow corresponding to the low power output The two combustors shown above hold nearly constant fuel-air ratios over the load range by